The term "composite" material would be very familiar to most engineers working in the aircraft or aerospace industries. There, composite materials are generally described as being materials that consist of fibers, such as graphite fibers, for example, embedded in a resinous matrix.
Because it is a lightweight metal, aluminum has traditionally been the material of choice for making the bulk of aircraft parts. However, since composites are even lighter, and have a better strength-to-weight ratio when compared to aluminum, aircraft designers have long been aware of the desirability of increasing the use of composites in aircraft structures. This has resulted in ongoing research and development activities directed to the development of new techniques for forming, curing and bonding composites into desired part shapes.
The "Stealth" aircraft developed by and for the military may provide the first and best example of large-scale composite usage for producing both aircraft fuselage and wing structures. However, it seems that composites were used there primarily because of their ability to reduce radar detection, as opposed to their superior material characteristics and what such characteristics could mean with respect to aircraft performance. Presently, there is an ongoing effort to produce large sections of conventional aircraft completely from composites. In particular, The Boeing Company, who is the assignee of the invention disclosed here, has been involved in producing a fuselage or "crown" panel section completely and solely from composite parts where the parts are bonded together as opposed to using other, more conventional fastening techniques.
It is relatively easy to produce composite skin panels, stiffeners or stringers, and circular frame members that define the curvature of a fuselage panel section. The problem lies in joining these various parts together as an intricate structure of criss-crossing frames and stringers, all of which are bonded to the skin panel.
Up to the present, producing such a structure is difficult at best, at least when conventionally-known methods for curing and bonding composites are used. Generally, conventional methods involve a high degree of labor, require non-reusable materials for forming and curing, and are very susceptible to cure failures.
Most or all composite parts are presently formed using a vacuum bag for pressing the uncured composite against a mandrel, and thereafter curing the composite at high temperature. Generally, some sort of intermediate materials must be positioned between the vacuum bag and the uncured composite, in order to avoid fouling the surface of the bag so that it is otherwise re-usable for forming other parts. With respect to the crown panel structure described above, positioning such protective materials between the bag and both the stringers and frame members is highly labor-intensive, and is not conducive to mass production or an automated process. Further, in order to reduce warpage and stress, it is desirable to co-cure the skin panel and stringers simultaneously. This creates technical problems with respect to the control of part shape as the stringers are pressed and cured against the skin panel, and with respect to controlling resin flow outwardly from the stringers.
It is these kinds of problems that the present invention addresses, and it is believed that the present invention makes the fabrication of intricate composite structures more practical from both a technical and economical standpoint.